Spacecraft are subject to heating by solar and infrared radiation which may affect their thermal and electrical performance. It is therefore desirable to know the expected absorbed heat flux in operating conditions. For this purpose, spacecraft are tested in thermal vacuum chambers to simulate orbital conditions and verify the thermal and electrical performance of the spacecraft.
The simulated orbital conditions can be achieved using the technique of IR or solar illumination. In either case it is desirable to measure accurately the radiated heat flux absorbed by the various spacecraft surfaces, and, for that reason, it has become necessary to design an accurate heat flux monitoring device or radiometer to measure, particularly, the absorbed IR radiated heat flux encountered during thermal vacuum testing.
Radiometers are difficult to design with the necessary accuracy. This is because it is difficult to isolate the heat flux sensing part of the radiometer, usually a disc or plate, so that the only heat transfer involving the sensing plate is that with the environment. The sensing plate must be fixed in place thus requiring a housing of some kind, and a temperature sensitive device, usually some kind of thermocouple, to be connected to the sensing plate. This causes conductive and radiated heat loss from the sensing plate to the thermocouple and housing. In addition the sensing plate radiates heat directly into the environment. Hence any measurement of radiated heat flux conducted by monitoring only the temperature of the sensing plate will be inaccurate.
Previous radiometers have sought to overcome this problem by, for example, thermally shielding the plate from the housing, attempting to minimize conductive and radiated losses (by using appropriate materials and coatings), and using calibration techniques with the housing maintained at a fixed temperature.